• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.1
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• pp.54-63
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• 2006

Ejector system are used to transport a low momentum flow to the higher pressure flow by the momentum change between high and low momentum flows. This system is used to simulate the high altitude and Mach number condition over altitude 20 km and Mach 4 of the supersonic test facility. We applied the design and the performance analysis technique(EISIMP code) of the Ramjet Test Facility(RJTF) air system in JAXA to the ejector system of the ramjet test facility in KARI. After preliminary design of the ejector system, we performed a computational study using FLUENT and investigated shock structures and flow characteristics of the ejector system.

• International Journal of Aerospace System Engineering
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• v.3 no.1
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• pp.25-29
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• 2016

This paper highlights the performance of an impulse turbine which is a part of turbopump in a liquid rocket first stage engine. The turbopump, currently under development at Korea Aerospace Research Institute, has an impulse type turbine with 12 nozzles and a single rotor. The impulse turbine can archive high specific power with the low gas flow rates. The supersonic impulse turbine with a single rotor can make a simple structure. High-pressure gases are converted into the dynamic energy with flows through the 12 nozzles and drive the rotor to make the power for the pumps. The turbine test was performed in the high-pressured turbine test facility with air gas instead of burned gas. A hydraulic dynamometer was used to absorb the power from the turbine and control the rotational speed and torque. The test points were at several pressure ratios with 7 different rotational speeds. Results showed the efficiency was highest at the design pressure ratio. The efficiency was insensitive to the pressure ratio variation than the rotational speed. It was a typical characteristic in an impulse turbine.

• Journal of the Korean Society of Propulsion Engineers
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• v.2 no.3
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• pp.10-19
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• 1998

When an under-expanded supersonic jet impinges on an inclined flat surface, a complex flow structure is established due to the intersection between the flat surface and the shock system of the free jet. This study reports on an experimental results of flows due to under-expanded axisymmetric sonic jets impinging on flat plate. Plate inclination from $60^{\cire}$~$90^{\cire}$ were investigated by means of detailed measurements of the surface pressure and schlieren photograph and surface flow visualization. The schlieren photograph are consistent with the pressure distribution and the surface flow visualization pictures are clearly related to the pressure distributions. The maximum wall pressure is found to be large on the inclined plate than on the perpendicular plate.

• Journal of the Korean Society of Propulsion Engineers
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• v.5 no.4
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• pp.50-56
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• 2001

The present study is an experimental work of the soni $c^ersonic air ejector-diffuser system. The pressure-time dependence in the secondary chamber of this ejector system is measured to investigate the steady operation of the ejector system. Six different primary nozzles of two sonic nozzles, two supersonic nozzles, petal nozzle, and lobed nozzle are employed to drive the ejector system at the conditions of different operating pressure ratios. Static pressures on the ejector-diffuser walls are to analyze the complicated flows occurring inside the system. The volume of the secondary chamber is changed to investigate the effect on the steady operation. the results obtained show that the volume of the secondary chamber does not affect the steady operation of the ejector-diffuser system but the time-dependent pressure in the secondary chamber is a strong function of the volume of the secondary chamber.er.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.192-197
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• 2004

Recently, fluidic thrust vectoring methods have been preferably employed to control the movement of propulsive systems due to relatively simpler design and lower cost than mechanical thrust vectoring methods. For An application of the thrust vectoring to flight bodies, it is necessary to understand very complicated exhaust flows which are often subject to shock waves and boundary layer separation. But researches for the thrust vector control using counterflow have been few. In the present study, experiments have been performed to investigate the characteristics of supersonic jets controlled by a thrust vectoring method using counterflow. The primary jet is expanded through a two-dimensional primary nozzle shrouded by collars, and is deflected by the suction of the air near nozzle into an upper slot placed between the primary nozzle and the upper collar. A shadowgraph method is used to visualize the supersonic jet flowfields. Primary nozzle pressure ratios and suction nozzle pressure ratios are varied from 3.0 to 5.0, and from 0.2 to 1.0 respectively. The present experimental results showed that, for a given primary nozzle pressure ratio, a decrease in the suction nozzle pressure ratio produced an increased thrust vector angle. As the suction nozzle pressure ratios were increased and decreased, the hysteresis of the thrust vectoring was observed through the wall pressure distributions

• Transactions of the Korean Society of Mechanical Engineers
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• v.13 no.3
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• pp.547-553
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• 1989

Last several stages of high capacity fossil power steam turbine and most stages of nuclear power steam turbine operate on wet steam. As a consequence, the flows in those cascades are accompanied by condensation, and the latent heat caused by condensation affects an oblique shock wave being generated at the vicinity of trailing of the blade. In the case of expanding of moist air through a suction type indraft wind tunnel, the effect of condensation affection the oblique shock wave generated by placing the small wedge into the supersonic part of the nozzle was investigated experimentally. In these connections, the relationship between condensation zone and reflection point of the incident oblique shock wave, angle between wedge bottom wall and oblique shock wave, and the variations of angles of incident and reflected shock waves due to the variation of initial stagnation relative humidity are discussed. Furthermore, the relationship between initial stagnation relative humidity and load working on the nozzle wall, obtained by measuring static pressure at the nozzle centerline, is discussed.

• Advances in aircraft and spacecraft science
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• v.7 no.3
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• pp.203-214
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• 2020

In many applications like the aircraft or the rockets/missiles, the flow from a nozzle needs to be expanded suddenly in an enlarged duct of larger diameter. The enlarged duct is provided after the nozzle to maximize the thrust created by the flow from the nozzle. When the fluid is suddenly expanded in an enlarged duct, the base pressure is generally lower than the atmospheric pressure, which results in base drag. The objective of this research work is to optimize the length to diameter (L/D) ratio of the enlarged duct using the CFD analysis in the flow field from the supersonic nozzle. The flow from the nozzle drained in an enlarged duct, the thrust, and the base pressure are studied. The Mach numbers for the study were 1.5, 2.0 and 2.5. The nozzle pressure ratios (NPR) of the study were 2, 5 and 8. The L/D ratios of the study were 0.5, 1, 2, 3, 4, 5, 6, 7, 8, 9 and 10. Based on the results, it is concluded that the L/D ratio should be increased to an optimum value to reattach the flow to an enlarged duct and to increase the thrust. The supersonic suddenly expanded flow field is wave dominant, and the results cannot be generalized. The optimized L/D ratios for various combinations of flow and geometrical parameters are given in the conclusion section.

• Journal of computational fluids engineering
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• v.18 no.2
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• pp.66-71
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• 2013

In this paper, the supersonic flows around space launch vehicles have been numerically simulated by using a 3-D RANS flow solver. The focus of the study was made for investigating plume-induced flow separation(PIFS). For this purpose, a vertex-centered finite-volume method was utilized in conjunction with 2nd-order Roe's FDS to discretize the inviscid fluxes. The viscous fluxes were computed based on central differencing. The Spalart-Allmaras model was employed for the closure of turbulence. The Gauss-Seidel iteration was used for time integration. To validate the flow solver, calculation was made for the 0.04 scale model of the Saturn-5 launch vehicle at the supersonic flow condition without exhaust plume, and the predicted results were compared with the experimental data. Good agreements were obtained between the present results and the experiment for the surface pressure coefficient and the Mach number distribution inside the boundary layer. Additional calculations were made for the real scale of the Saturn-5 configuration with exhaust plume. The flow characteristics were analyzed, and the PIFS distances were validated by comparing with the flight data. The KSLV-1 is also simulated at the several altitude conditions. In case of the KSLV-1, PIFS was not observed at all conditions, and it is expected that PIFS is affected by the nozzle position.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.1
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• pp.52-60
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• 2013

Ejector-diffuser system makes use of high-pressure primary stream to entrain the low-pressure secondary stream through pure shear action between two streams. In general, the flow field in the ejector-diffuser system is highly complicated due to turbulent mixing, compressibility effects. A fatal drawback of the ejector system is in its low efficiency. Many works have been done to improve the performance of the ejector system, but not yet satisfactory. In the present study, a mixing guide vane was installed at the inlet of the secondary stream for the purpose of the performance improvement of the ejector system. A CFD method has been applied to simulate the supersonic flows inside the ejector-diffuser system. The present results obtained were validated with existing experimental data. The mixing guide vane effects are discussed in terms of the entrainment ratio, total pressure loss as well as pressure recovery.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.30 no.6
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• pp.53-60
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• 2002

Based on the analyses of the single transverse injection in supersonic flow fields, the mixing characteristics of dual transverse injection of hydrogen in supersonic air flow are studied with computational methods. Three-dimensional Navier -Stokes and the k-$\omega$ SST turbulence model were used. A parametric study is conducted with the variation of the distance between two injectors. The flow patterns and the mixing characteristics of two injection flows are very different from each other, and the flow patterns and the mixing characteristics of the rear injection flow are strongly influenced by those of the first injection flow. The increase of the distance between two injectors up to a specific distance results in the increase of mixing rate and penetration of fuel. However, the increase of the distance over the specific distance results in the decrease of mixing rate and penetration of fuel. From the results it can be stated that there exists a distance between two injectors for optimum mixing characteristics.